This invention generally relates to a gas turbine engine, and more particularly to a nacelle for a turbofan gas turbine engine.
In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages that extract energy from the gases. In a two spool gas turbine engine, a high pressure turbine powers the high pressure compressor, while a low pressure turbine powers a fan disposed upstream of the compressor and a low pressure compressor.
Combustion gases are discharged from the turbofan engine through a core exhaust nozzle, and fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust provided from the combustion gases is discharged through the core exhaust nozzle.
In high bypass turbofans a majority of the air pressurized by the fan bypasses the turbofan engine for generating propulsion thrust. High bypass turbofans typically use large diameter fans to achieve adequate turbofan engine efficiency. Therefore, the nacelle of the turbofan engine must be large enough to support the large diameter fan of the turbofan engine. Disadvantageously, the relatively large size of the nacelle results in increased weight and drag that may offset the propulsive efficiency achieved by high bypass turbofan engines.
Accordingly, it is desirable to optimize the performance of a gas turbine engine during diverse flight requirements to provide a nacelle having a reduced maximum diameter, reduced weight, and reduced drag.